1. Field of the Invention
The present invention relates to gas turbine engines and, more particularly, to blade and disk interfaces of such engines.
2. Background Art
Fan rotors can be manufactured integrally or as an assembly of blades around a disk. In the case where the rotor is assembled, the fixation between each blade and the disk has to provide retention against extremely high radial loads. This in turn causes high radial stress in the disk retaining the blades.
In the case of “swept” fans, the blades are asymmetric with respect to their radial axis. A significant portion of the weight of these blades is cantilevered over the front portion of the fixation, which causes an uneven axial distribution of the radial load on the fixation and disk. This load distribution causes high local radial stress in the front of the disk and high contact forces between the blade and the front of the disk.
Although a number of solutions have been provided to even axial distribution of stress in blades, such as grooves in blade platforms to alleviate thermal and/or mechanical stresses, these solutions do not address the problem of high local radial stress in the disk supporting the blades.
Some solutions have also been provided to reduce the increase of contact stress resulting in a non-zero broach angle of the blade, including the elimination of diagonally opposite portions of the load transfer interface which are less stressed. However, such solutions are not applicable to reduce the increased local contact stress produced by the asymmetry of “swept” fans. In addition, such solutions do not address the problem of high local radial stress in the disk supporting the blades.
Accordingly, there is a need for a blade and disk interface for a gas turbine engine fan producing reduced local contact stress and reduced local radial stress in the disk.